WO2016007116A1 - Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz - Google Patents

Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz Download PDF

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Publication number
WO2016007116A1
WO2016007116A1 PCT/US2014/045512 US2014045512W WO2016007116A1 WO 2016007116 A1 WO2016007116 A1 WO 2016007116A1 US 2014045512 W US2014045512 W US 2014045512W WO 2016007116 A1 WO2016007116 A1 WO 2016007116A1
Authority
WO
WIPO (PCT)
Prior art keywords
tip
rail
fin
slot
downstream
Prior art date
Application number
PCT/US2014/045512
Other languages
English (en)
Inventor
Ching-Pang Lee
Kok-Mun Tham
Gm Salam Azad
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to US15/318,001 priority Critical patent/US9810074B2/en
Priority to PCT/US2014/045512 priority patent/WO2016007116A1/fr
Priority to JP2017501032A priority patent/JP6347892B2/ja
Priority to EP14742442.8A priority patent/EP3167161A1/fr
Priority to CN201480080443.4A priority patent/CN106471215B/zh
Publication of WO2016007116A1 publication Critical patent/WO2016007116A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/50—Building or constructing in particular ways
    • F05D2230/53—Building or constructing in particular ways by integrally manufacturing a component, e.g. by milling from a billet or one piece construction
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the invention relates to gas turbine engine blade squealer tips and methods for cooling gas turbine engine squealer tips. More particularly, embodiments of the
  • Segmented suction side rail embodiments abrade opposing turbine casing abradable surfaces prior to potential contact with the pressure side rail, reducing likelihood of pressure side rail friction heating.
  • known turbine engines such as the gas turbine engine 30 include a multi stage compressor section 32, a combustor section 34, a multi stage turbine section 36 and an exhaust system 38. Atmospheric pressure intake air is drawn into the compressor section 32 generally in the direction of the flow arrows F along the axial length of the turbine engine 30. The intake air is progressively
  • the engine's rotor and shaft 39 has a plurality of rows of airfoil cross sectional shaped turbine blades 40 terminating in distal blade squealer tips 46 in the compressor 32 and turbine 36 sections.
  • Each blade 40 has a concave profile pressure side 42 and a convex suction side 44.
  • the high temperature and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on the blades 40, spinning the rotor 39s.
  • some of the mechanical power imparted on the rotor shaft is available for performing useful work.
  • the combustion gasses are constrained radially distal the rotor by turbine casing 60 and proximal the rotor by air seals.
  • respective upstream vanes 62 direct upstream combustion gases generally parallel to the incident angle of the leading edge 48 of turbine blade and downstream vanes redirect downstream combustion gas exiting the trailing edge 50 of the blade.
  • the turbine engine 30 turbine casing 60 proximal the blade squealer tips 46 is lined with a plurality of sector shaped abradable components 64, each having a support surface retained within and coupled to the casing 60 and an abradable substrate 66 that is in opposed, spaced relationship with the blade tip by a blade tip gap G.
  • the abradable substrate is often constructed of a metallic/ceramic material that has high thermal and thermal erosion resistance and that maintains structural integrity at high combustion temperatures.
  • metallic-ceramic materials is often more abrasive than the turbine blade tip 46 material a blade tip gap G is maintained to avoid contact between the two opposed components that might at best cause premature blade tip wear and in worse case circumstances might cause engine damage.
  • each respective blade tip 46 desirably has a uniform blade tip gap G relative to the abradable component 64 that is as small as possible (ideally zero clearance) to minimize blade tip airflow leakage L between the concave pressure blade side 42 and the convex suction blade side 44 as well as axially in the combustion flow direction F.
  • manufacturing and operational tradeoffs require blade tip gaps G greater than zero.
  • Such tradeoffs include tolerance stacking of interacting components, so that a blade constructed on the higher end of acceptable radial length tolerance and an abradable component abradable substrate 66 constructed on the lower end of acceptable radial tolerance do not impact each other excessively during operation.
  • small mechanical alignment variances during engine assembly can cause local variations in the blade tip gap G.
  • very small mechanical alignment variances can impart local blade tip gap G variances of a few millimeters.
  • the turbine engine casing 60 may experience out of round (e.g., egg shaped) thermal distortion.
  • Casing 60 thermal distortion potential increases between operational cycles of the turbine engine 30 as the engine is fired up to generate power and subsequently cooled for servicing after thousands of hours of power generation.
  • greater casing 60 and abradable component 64 distortion tends to occur at the uppermost and lowermost casing circumferential positions (i.e., 6:00 and 12:00 positions) compared to the lateral right and left circumferential positions (i.e., 3:00 and 9:00).
  • one or more of the blade tip squealers 46 may be worn during operation, increasing the blade tip gap locally in various other less deformed circumferential portions of the turbine casing 60 from the ideal gap G to a larger gap.
  • the excessive blade gap distortion increases blade tip leakage L, diverting hot combustion gas away from the turbine blade 40 airfoil, reducing the turbine engine's efficiency.
  • the exemplary blade 40 squealer tip 46 construction and its interaction with the turbine casing abradable surface 66 is shown in greater detail in FIGs. 3-6.
  • the squealer tip 46 has a an airfoil planform tip plate 56 having along its outer periphery downstream from its leading edge 48 and upstream from its trailing edge 50 opposed and laterally separated outwardly or radially projecting concave pressure 52 and convex suction 54 rails, which respectively have opposed inner faces and outer faces.
  • An enclosed tip cavity 57 is defined between the tip plate 56 and respective inner faces of the pressure rail 52 (also referenced in FIG. 4 as the pressure rail inner surface 53) and suction rail 54 from the leading 48 to trailing 50 edges.
  • pressure side gas flow F P is deflected around the leading edge 48 and separates from contact with the pressure side rail 52, allowing heat to concentrate on the outer face of the pressure rail.
  • Such excessive heat concentration can cause pressure rail 52 erosion, prematurely wearing out the blade and undesirably increasing the blade tip gap, as previously described.
  • Combustion gas flow F T undesirably passes through the blade tip gap over the top of the squealer tip 46, but most of it is diverted away from the pressure rail inner surface 53 toward the suction side rail, creating another potential heat concentration zone along the pressure rail inner surface.
  • Gas flow F s along the suction side 44 of the blade tip 46 is directed toward the blade trailing edge 50, where it cannot assist in transfer of heat from the pressure rail 52 heat concentration zone.
  • friction contact between the squealer tip 46 pressure rail 52 and the abradable surface 46 also undesirably increases pressure rail area heat concentration.
  • FIG. 7 Another known conventional blade squealer tip 146 is shown in FIG. 7, having a segmented pressure side rail 152 with a slot 158 proximal the squealer tip 146 trailing edge 150.
  • the suction side rail 154 is continuous downstream from the leading edge 148 to the trailing edge 150.
  • the rails 152, 154 and the underlying tip plate (not shown) form the squealer tip cavity 157.
  • a suggested object is to reduce turbine blade squealer tip wear by decreasing squealer tip pressure rail operating temperature through increased cooling air flow along an inside surface of the pressure rail.
  • Another suggested object is to reduce turbine blade squealer tip wear by decreasing squealer tip pressure rail operating temperature through reduced contact between the pressure rail and the engine's opposed abradable surface. Reduction or elimination of pressure rail contact with the abradable surface reduces likelihood of rubbing friction heating of the pressure side rail.
  • gas turbine engine blade squealer tips that incorporate cooling slots formed in the suction side rail downstream of the leading edge for directing cooling gas flow along an inside edge of the squealer tip pressure side rail.
  • Some embodiments incorporate a tip fin on the suction side rail proximal a cooling slot.
  • Segmented suction side rail embodiments abrade opposing turbine casing abradable surfaces prior to potential contact with the pressure side rail, reducing likelihood of pressure side rail friction heating.
  • cooler pressure side rails reduce likelihood of squealer tip erosion.
  • Exemplary embodiments feature a gas turbine engine blade squealer tip, comprising an airfoil planform tip plate having along its outer periphery downstream from its leading edge and upstream from its trailing edge opposed and laterally separated projecting concave pressure and convex suction rails respectively having inner and outer faces.
  • An enclosed tip cavity is defined between the tip plate and respective inner faces of the pressure and suction rails from the leading to trailing edges.
  • At least one slot is formed through respective inner and outer faces of the suction rail downstream of the leading edge. The slot is in communication with the tip cavity and is oriented for directing cooling air flow there through and downstream along the pressure rail inner face.
  • blade squealer tips in method embodiments for cooling a gas turbine engine that includes a rotor having blades radially projecting therefrom, with blade squealer tips in opposed relationship with a circumferential abradable layer supported by a turbine casing.
  • the method is performed by providing and installing turbine blades having the afore described blade squealer tips and operating the engine so that cooling air flows downstream along the pressure rail inner face and through the slot that is formed through respective inner and outer faces of the suction rail downstream of the leading edge.
  • Additional embodiments feature a method for manufacturing a gas turbine engine blade squealer tip pressure side rail by providing a turbine blade with an airfoil planform tip plate having along its outer periphery downstream from its leading edge and upstream from its trailing edge opposed and laterally separated projecting concave pressure and convex suction rails respectively having inner and outer faces and an enclosed tip cavity defined between the tip plate and respective inner faces of the pressure and suction rails from the leading to trailing edges.
  • a location is determined for at least one slot in the blade tip through respective inner and outer faces of the suction rail downstream of the leading, with the slot in communication with the tip cavity and oriented for directing cooling air flow there through and downstream along the pressure rail inner face.
  • the slot is formed in the blade tip at the determined location.
  • FIG. 1 Other embodiments feature a gas turbine engine, comprising a rotor having blades radially projecting therefrom, with each blade having a squealer tip including an airfoil planform tip plate having along its outer periphery downstream from its leading edge and upstream from its trailing edge opposed and laterally separated projecting concave pressure and convex suction rails respectively having inner and outer faces.
  • the squealer tip includes an enclosed tip cavity defined between the tip plate and respective inner faces of the pressure and suction rails from the leading to trailing edges.
  • At least one slot is formed through respective inner and outer faces of the pressure rail downstream of the leading edge. Each respective slot is in communication with the tip cavity and is oriented for directing cooling air flow there through and downstream along the pressure rail inner face.
  • FIG. 1 is a partial axial cross sectional view of an exemplary known gas turbine engine
  • FIG. 2 is a detailed cross sectional elevational view of a known Row 1 turbine blade and vanes showing blade tip gap G between a blade tip and abradable component of the turbine engine of FIG. 1;
  • FIG. 3 is a perspective view of the exemplary known turbine blade of FIGs. 1 and
  • FIG. 4 is an elevational cross sectional view of the known turbine blade and squealer tip of FIG. 3 taken along 3-3;
  • FIG. 5 is a schematic plan form view of the known squealer tip of FIGs. 3 and 4 its opposed orientation and motion relative to a turbine engine abradable surface;
  • FIG. 6 is a streamline flow simulation of gas flow around the known turbine blade squealer tip and abradable surface of FIG. 5;
  • FIG. 7 is a schematic plan form view similar to FIG. 5, of another known squealer tip and its opposed relative orientation and motion relative to a turbine engine abradable surface;
  • FIG. 8 is a schematic plan form view similar to FIG. 7, of an exemplary first embodiment of a squealer tip of the invention and its opposed relative orientation and motion relative to a turbine engine abradable surface
  • FIG. 9 is a schematic plan form view similar to FIG. 7, of an exemplary second embodiment of a squealer tip of the invention and its opposed relative orientation and motion relative to a turbine engine abradable surface;
  • FIG. 10 is a top elevational view of a turbine blade that incorporates the first embodiment squealer tip of FIG. 8;
  • FIG. 11 is a perspective view of the turbine blade of FIG. 10;
  • FIG. 12 is a streamline flow simulation of gas flow around the turbine blade with the first embodiment squealer tip of FIG. 8;
  • FIG. 13 is a top elevational view of a turbine blade that incorporates the second embodiment squealer tip of FIG. 9;
  • FIG. 14 is a perspective view of the turbine blade of FIG. 13;
  • FIG. 15 is a streamline flow simulation of gas flow around the turbine blade with the second embodiment squealer tip of FIG. 9.
  • turbine blade squealer tips incorporate one or more cooling slots formed in the suction side rail downstream of the leading edge. These slots are oriented for directing cooling gas flow along an inside edge of the squealer tip pressure side rail, so that heat concentration along the pressure side rail is transported away from hottest zone of the squealer tip.
  • Some embodiments incorporate a tip fin on the suction side rail proximal a cooling slot.
  • Segmented suction side rail embodiments abrade opposing turbine casing abradable surfaces (analogous to a snow plow) prior to potential contact with the pressure side rail, reducing likelihood of pressure side rail friction heating.
  • cooler pressure side rails reduce likelihood of squealer tip erosion.
  • the known conventional blade tip 46/146 has a unified, continuous squealer rail of uniform thickness on both concave pressure 52/152 and convex suction 54/154 sides.
  • the suction side squealer is the first to cut into the ring segment. From the gas flow simulation CFD analysis as shown in FIG.
  • the gas flow past the leading edge 48 of the tip 46 splits into two streams, one toward the pressure side 42 and one toward the suction side 44.
  • the suction side gas stream Fs enters into the tip cavity at the forward section and mixes with the leakage flow from the pressure side Fp at the downstream location before exiting to the suction side in the downstream section.
  • the invention embodiments of FIGs. 8 and 9, each respectively with a segmented suction side squealer 254/354 allows more of the suction side gas stream F s to enter into the tip cavity 257/357 and pressurize the tip cavity (analogous to a static wall) that will lead to less leakage Fp from the pressure side 252/352.
  • the segmented squealer designs that include the
  • fins 262/264/254 or 364/354 provide laterally overlapped squealers on their respective suction side to have more cutting power to the abradable ring segment patterns and have a better chance to preserve the pressure side squealer 252/352 for better sealing.
  • the segmented and overlapped suction side 262/264/254 or 364/354 squealer construction embodiments of FIGs. 8 and 9 have more durable blade tips and less performance robbing tip leakage than the conventional squealer tip 46/146 designs of FIGs. 5 and 7.
  • Two exemplary embodiments of squealer tips constructed in accordance with the teachings of the invention are shown in FIGs. 8-15.
  • a first exemplary embodiment blade 240 with squealer tip 246 is shown in FIGs. 8 and 10-12, having the previously described segmented suction side downstream of the leading edge 248, formed from first fin 262, second fin 264 and suction rail 254.
  • First slot 260 and second slot 266 allow communication between the suction side of the blade 240 and the tip cavity 257, as does the optional slot 258 formed in the pressure rail 252 proximal the trailing edge 250.
  • the squealer tip is formed with the first and second slots 260, 266 with or without the slot 258.
  • cooling gas flow F T within the cavity 257 is directed along the pressure rail inner face 253, thereby transporting heat away from the pressure rail 252.
  • FIGs. 9 and 13-15 A second exemplary embodiment blade 340 with squealer tip 346 is shown in FIGs. 9 and 13-15, having the previously described segmented suction side downstream of the leading edge 348, formed from first fin 362 and suction rail 354.
  • First slot 360 allows communication between the suction side of the blade 340 and the tip cavity 357, as does the optional slot 358 formed in the pressure rail 352 proximal the trailing edge 350.
  • the squealer tip 346 is formed with the first slot 360 with or without the slot 358.
  • cooling gas flow F T within the cavity 357 is directed along the pressure rail inner face 353, thereby transporting heat away from the pressure rail 352.
  • Additional beneficial gas flow through the squealer tip cavity 357 along the pressure rail inner face 353 is optionally provided by adding cooling holes 370 along the suction side or cooling holes 372 in the tip cavity or at both locations.
  • connection means “connected,” “supported,” and “coupled” and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further,

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne des bouts amincis d'aube de turbine à gaz qui renferment des fentes de refroidissement formées dans le rail latéral d'aspiration en aval du bord d'attaque pour diriger un écoulement de gaz de refroidissement le long d'un bord intérieur du rail latéral de pression à bout aminci. Certains modes de réalisation comprennent une ailette à pointe sur le rail latéral d'aspiration à proximité d'une fente de refroidissement. Les modes de réalisation de rail latéral d'aspiration segmenté prévoient l'abrasion de surfaces abradables opposées d'un carter de turbine avant un contact potentiel avec le rail latéral de pression, réduisant la probabilité d'un échauffement par frottement du rail latéral de pression. Pendant le fonctionnement de la turbine, des rails latéraux de pression de refroidissement réduisent les risques d'érosion de l'aube à bout aminci.
PCT/US2014/045512 2014-07-07 2014-07-07 Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz WO2016007116A1 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US15/318,001 US9810074B2 (en) 2014-07-07 2014-07-07 Segmented turbine blade squealer tip and cooling method
PCT/US2014/045512 WO2016007116A1 (fr) 2014-07-07 2014-07-07 Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz
JP2017501032A JP6347892B2 (ja) 2014-07-07 2014-07-07 ガスタービンエンジンブレードスキーラ先端、対応する製造および冷却方法、ならびにガスタービンエンジン
EP14742442.8A EP3167161A1 (fr) 2014-07-07 2014-07-07 Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz
CN201480080443.4A CN106471215B (zh) 2014-07-07 2014-07-07 燃气涡轮叶片凹槽状叶顶、对应的制造和冷却方法及燃气涡轮发动机

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2014/045512 WO2016007116A1 (fr) 2014-07-07 2014-07-07 Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz

Publications (1)

Publication Number Publication Date
WO2016007116A1 true WO2016007116A1 (fr) 2016-01-14

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PCT/US2014/045512 WO2016007116A1 (fr) 2014-07-07 2014-07-07 Bout aminci d'aube de turbine à gaz, procédés de fabrication et de refroidissement correspondants et turbine à gaz

Country Status (5)

Country Link
US (1) US9810074B2 (fr)
EP (1) EP3167161A1 (fr)
JP (1) JP6347892B2 (fr)
CN (1) CN106471215B (fr)
WO (1) WO2016007116A1 (fr)

Cited By (3)

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US20160319672A1 (en) * 2015-04-29 2016-11-03 General Electric Company Rotor blade having a flared tip
US20180328191A1 (en) * 2017-05-10 2018-11-15 General Electric Company Rotor blade tip
US10533429B2 (en) * 2017-02-27 2020-01-14 Rolls-Royce Corporation Tip structure for a turbine blade with pressure side and suction side rails

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Publication number Priority date Publication date Assignee Title
CN107035844B (zh) * 2017-05-25 2021-02-02 吉林大学 一种液力变矩器分段式涡轮叶片
US10808572B2 (en) 2018-04-02 2020-10-20 General Electric Company Cooling structure for a turbomachinery component
US11118462B2 (en) * 2019-01-24 2021-09-14 Pratt & Whitney Canada Corp. Blade tip pocket rib
US11371359B2 (en) 2020-11-26 2022-06-28 Pratt & Whitney Canada Corp. Turbine blade for a gas turbine engine

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US4761116A (en) * 1987-05-11 1988-08-02 General Electric Company Turbine blade with tip vent
EP1057972A2 (fr) * 1999-06-01 2000-12-06 General Electric Company Joint d'extrémité d'aube de turbine décalé
US20120237358A1 (en) * 2011-03-17 2012-09-20 Campbell Christian X Turbine blade tip
US20140037458A1 (en) * 2012-08-03 2014-02-06 General Electric Company Cooling structures for turbine rotor blade tips

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US5503527A (en) * 1994-12-19 1996-04-02 General Electric Company Turbine blade having tip slot
US6059530A (en) * 1998-12-21 2000-05-09 General Electric Company Twin rib turbine blade
EP2351908B1 (fr) * 2008-10-30 2016-08-17 Mitsubishi Hitachi Power Systems, Ltd. Aube rotorique de turbine
US20130149163A1 (en) * 2011-12-13 2013-06-13 United Technologies Corporation Method for Reducing Stress on Blade Tips

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Publication number Priority date Publication date Assignee Title
US4761116A (en) * 1987-05-11 1988-08-02 General Electric Company Turbine blade with tip vent
EP1057972A2 (fr) * 1999-06-01 2000-12-06 General Electric Company Joint d'extrémité d'aube de turbine décalé
US20120237358A1 (en) * 2011-03-17 2012-09-20 Campbell Christian X Turbine blade tip
US20140037458A1 (en) * 2012-08-03 2014-02-06 General Electric Company Cooling structures for turbine rotor blade tips

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20160319672A1 (en) * 2015-04-29 2016-11-03 General Electric Company Rotor blade having a flared tip
US10533429B2 (en) * 2017-02-27 2020-01-14 Rolls-Royce Corporation Tip structure for a turbine blade with pressure side and suction side rails
US20180328191A1 (en) * 2017-05-10 2018-11-15 General Electric Company Rotor blade tip
US10443405B2 (en) * 2017-05-10 2019-10-15 General Electric Company Rotor blade tip

Also Published As

Publication number Publication date
US9810074B2 (en) 2017-11-07
JP2017529476A (ja) 2017-10-05
CN106471215A (zh) 2017-03-01
JP6347892B2 (ja) 2018-06-27
US20170122110A1 (en) 2017-05-04
CN106471215B (zh) 2018-06-19
EP3167161A1 (fr) 2017-05-17

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